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98-26-16: This amendment adopts a new airworthiness directive (AD) that applies to certain Raytheon Aircraft Company (Raytheon) Models 1900, 1900C, and 1900D airplanes. This AD requires modifying the emergency exit doors and installing interior and exterior placards on each of the emergency exit doors. Difficulty in opening the emergency exit doors prompted this action. The actions specified by this AD are intended to prevent passengers and crew from not being able to open the emergency exit doors during an airplane emergency, which could result in passenger and crew injuries.
2013-13-15: We are superseding airworthiness directive (AD) 87-02-07, which applied to all The Boeing Company Model 737-100 and -200 series airplanes. AD 87-02-07 required replacement of certain underwing fuel tank access covers with stronger, fire-resistant covers. This new AD also requires inspecting fuel tank access doors to determine that impact-resistant access doors are installed in the correct locations, inspecting application of stencils and index markers of impact- resistant access doors, doing corrective actions if necessary, revising the maintenance program, and adding airplanes to the applicability. This AD was prompted by reports of standard access doors installed where impact-resistant access doors are required, and reports of impact-resistant doors without stencils. We are issuing this AD to prevent foreign object penetration of the wing tank, which could lead to a fuel leak near ignition sources (engine, hot brakes), consequently leading to a fuel-fed fire.
93-01-13: 93-01-13 FOKKER: Amendment 39-8467. Docket 92-NM-236-AD. Applicability: Model F27 series airplanes; except Model F27 Mark 050 series airplanes; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent uncontrolled fuel leakage into the wheel well, accomplish the following: (a) Within 30 days after the effective date of this AD, perform a visual inspection to detect chafing, scratches, or dents on the fuel lines on the left- and right-hand nacelles, in accordance with Part 1 of the Accomplishment Instructions of Fokker Service Bulletin F27/28-61, dated September 3, 1992. (b) If no chafing, scratches, or dents are found during the inspection required by paragraph (a) of this AD, the fuel line may remain installed. However, within 100 hours time-in-service after the effective date of this AD, perform a visual inspection for clearances between the fuel line and the landing gear parts, in accordance with Part2 of the Accomplishment Instructions of Fokker Service Bulletin F27/28-61, dated September 3, 1992; and accomplish either paragraph (b)(1) or (b)(2) of this AD, as applicable. (1) If the clearance between the fuel line and the landing gear parts is less than 0.25 inch, prior to further flight, adjust the landing gear components to achieve a minimum clearance of 0.25 inch, in accordance with the service bulletin. (2) If the clearance between the fuel line and the landing gear parts is 0.25 inch or greater, no further action is required by this AD. (c) If any chafing or scratches are found during the inspection required by paragraph (a) of this AD, prior to further flight, accomplish both paragraphs (c)(1) and (c)(2) of this AD: (1) Measure the damage depth and accomplish either paragraph (c)(1)(i), (c)(1)(ii,) or (c)(1)(iii) of this AD, as applicable. (i) If the depth of chafing or scratch damage is equal to or greater than 0.009 inch, prior to further flight, replace the damaged fuel line, in accordance with Chapter 28-00-00 of the Airplane Maintenance Manual. (ii) If the depth of chafing or scratch damage is 0.004 inch or more, but less than 0.009 inch, re-inspect the damaged fuel line in accordance with paragraph (a) of this AD thereafter at intervals not to exceed 10 hours time-in-service; and replace the damaged fuel line within 50 hours time-in- service after the effective date of this AD, in accordance with Chapter 28-00-00 of the Airplane Maintenance Manual. (iii) If the depth of chafing or scratch damage is less than 0.004 inch, the fuel line may remain in place. (2) Perform a visual inspection for clearances between the fuel line and the landing gear parts, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/28-61, dated September 3, 1992, and accomplish either paragraph (c)(2)(i) or (c)(2)(ii) of this AD, as applicable. (i) If the clearance between the fuel lineand the landing gear parts is less than 0.25 inch, prior to further flight, adjust the landing gear components to achieve a minimum clearance of 0.25 inch, in accordance with the service bulletin. (ii) If the clearance between the fuel line and the landing gear parts is 0.25 inch or greater, no further action is required by this AD. (d) If any dent is found during the inspection required by paragraph (a) of this AD, prior to further flight, accomplish both paragraphs (d)(1) and (d)(2) of this AD: (1) Measure the damage depth and accomplish either paragraph (d)(1)(i) or (d)(1)(ii) of this AD, as applicable. (i) If the depth of the dent is equal to or greater than 0.1 inch, prior to further flight, replace the dented fuel line, in accordance with Chapter 28-00-00 of the Airplane Maintenance Manual. (ii) If the depth of the dent is less than 0.1 inch, the fuel line may remain installed. (2) Perform a visual inspection for clearances between the fuel line and the landing gear parts, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/28-61, dated September 3, 1992; and accomplish either paragraph (d)(2)(i) or (d)(2)(ii) of this AD, as applicable. (i) If the clearance between the fuel line and the landing gear parts is less than 0.25 inch, prior to further flight, adjust the landing gear components to achieve a minimum clearance of 0.25 inch, in accordance with the service bulletin. (ii) If the clearance between the fuel line and the landing gear parts is 0.25 inch or greater, no further action is required by this AD. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then sendit to the Manager, Standardization Branch. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch. (f) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (g) The inspections, replacement, and adjustment shall be done in accordance with Fokker Service Bulletin F27/28-61, dated September 3, 1992. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC.(h) This amendment becomes effective on February 5, 1993.
91-05-02: 91-05-02 SOCATA: Amendment 39-6887. Docket No. 90-CE-20-AD. Applicability: Models TB 20 and TB 21 airplanes (serial numbers (S/N) 1 through 1051, except S/N 1040 and S/N 1042), certificated in any category. Compliance: Required as indicated after the effective date of this AD, unless already accomplished. To prevent structural failure of the fuselage frame in the area of the landing gear attachment, accomplish the following: (a) On airplanes with more than 1,500 hours time-in-service (TIS) on the effective date of this AD, within the next 100 hours TIS after the effective date of this AD and, thereafter, at intervals not to exceed 500 hours TIS, visually inspect the fuselage frame No. 0 for cracks in the area of the engine mount and landing gear mount in accordance with the instructions in Aerospatiale Service Bulletin (SB) No. 42/1, dated July 1990. Prior to further flight, repair any cracked frames found in accordance with the instructions in the above SB.(b) On airplanes with less than 1,500 hours TIS on the effective date of this AD, within the next 100 hours TIS or prior to accumulating 1,600 hours TIS, whichever occurs later, and, thereafter, at intervals not to exceed 500 hours TIS, visually inspect the fuselage frame No. 0 for cracks in the area of the engine mount and landing gear mount in accordance with the instructions in Aerospatiale SB No. 42/1, dated July 1990. Prior to further flight repair any cracked frames found per the instructions in the above SB. (c) The repetitive inspections specified in paragraphs (a) and (b) of this AD are no longer required when the airplane has been modified in accordance with Socata Kit 9152. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. (e) An alternate method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Aircraft Certification Office, Europe, Africa, Middle East Office, FAA, c/o American Embassy, B-1000, Brussels, Belgium. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. (f) All persons affected by this directive may obtain copies of the documents referred to herein upon request to Aerospatiale Aeroport Tarbes-Ossum-Lourdes, B.P. 930 65009 Tarbes, France; Telephone 62.51.7300; or may examine the service information at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This amendment (39-6887, AD 91-05-02) becomes effective on March 25, 1991.
98-26-15: This amendment supersedes an existing airworthiness directive (AD), applicable to certain British Aerospace (Jetstream) Model 4101 airplanes, that currently requires repetitive detailed visual inspections to detect cracks in the shear cleats of the roller guide structural support of the passenger door, and replacement of any cracked shear cleat with a new shear cleat. That AD also provides for an optional terminating modification that constitutes terminating action for the repetitive inspections. This amendment mandates accomplishment of the previously optional terminating modification. This amendment is prompted by reports indicating that fatigue cracking was detected in the roller guide shear cleats of the passenger door. The actions specified by this AD are intended to prevent such fatigue-related cracking, which could result in structural failure or loss of the passenger door, and consequent rapid depressurization of the airplane during flight.
97-21-10: 97-21-10 AIRBUS INDUSTRIE: Amendment 39-10163. Docket 97-NM-265-AD. Applicability: Model A319, A320, and A321 series airplanes, certificated in any category; on which any of the following Airbus Modifications have been installed: Affected Model(s) Airbus Modification Installed A319 and A321 A319, A320, and A321 25469 (reference Airbus Service Bulletin A320-22-1054) 26093 A320 24065 (reference Airbus Service Bulletin A320-22-1040) or 24067 (reference Airbus Service Bulletin A320-22-1039) A320 25314 (reference Airbus Service Bulletin A320-22-1051) or 25315 (reference Airbus Service Bulletin A320-22-1050) A320 and A321 24064 (reference Airbus Service Bulletin A320-22-1034) or 24066 (reference Airbus Service Bulletin A320-22-1029) A320 and A321 25199 (reference Airbus Service Bulletin A320-22-1045) or 25200 (reference Airbus Service Bulletin A320-22-1046) A320 and A321 25240 (reference Airbus Service Bulletin A320-22-1033) or 25274 (reference AirbusService Bulletin A320-22-1056) A319, A320, and A3 26243 NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (b) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. Compliance: Required as indicated, unless accomplished previously. To ensure that the flightcrew detects and corrects an unintended flight path if certain software anomalies of the FMGS occur, which could result in an increased risk of collision with terrain or other airplanes, accomplish the following: (a) Within 10 days after the effective date of this AD, revise the Normal Procedures Section of the FAA-approved Airplane Flight Manual (AFM) by inserting a copy of Model A319/320/321 Flight Manual Temporary Revision 4.03.00/02, dated May 28, 1997, into the AFM. NOTE 2: When the temporary revision specified in paragraph (a) of this AD has been incorporated into the general revisions of the AFM, the general revisions may be inserted in the AFM, provided the information contained in the general revisions is identical to that specified in Model A319/320/321 Flight Manual Temporary Revision 4.03.00/02. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Operations Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE 3: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (c) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The AFM revision shall be done in accordance with Model A319/320/321 Flight Manual Temporary Revision 4.03.00/02, dated May 28, 1997. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from Airbus Industrie, 1 Rond Point Maurice Bellonte, 31707 Blagnac Cedex, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. NOTE 4: The subject of this AD is addressed in French airworthiness directive 97-153- 100(B), dated July 16, 1997. (e) This amendment becomes effective on November 3, 1997.
60-01-04: 60-01-04 FAIRCHILD: Amdt. 81 Part 507 Federal Register January 16, 1960. Applies to all Models F-27, F-27A, and -27B aircraft. Compliance required by January 15, 1960. Cases of "rudder walk" have been experienced on aircraft in service. Such rudder oscillation creates a flight hazard. In order to correct or prevent this condition, unless already accomplished, a beaded angle should be added to the rudder balance tab trailing edge in accordance with Fairchild Service Bulletin 27-14.
77-04-05: 77-04-05 CESSNA: Amendment 39-2838. Applies to 180 Series (Serial Numbers 18050662 thru 18052711); 182 Series (Serial Numbers 18251557 thru 18264790); 185 Series (Serial Numbers 18500001 thru 18501832); 188 Series (Serial Numbers 18800001 thru 18802348); and A188 Series (Serial Numbers 18800001 thru 18800707) airplanes. Compliance: Required as indicated, unless already accomplished. To prevent ingestion of the induction air box seal into the carburetor, within the next 50 hours' time in service after the effective date of this AD, accomplish the following: A. Disconnect the carburetor air intake flexible duct assembly from the induction air box. B. Visually inspect the condition of the existing carburetor induction air box seal(s) and if not securely bonded to the carburetor intake flexible duct assembly, remove the existing seal(s) and install new P/N 0752016-10 and -11 seals in accordance with the following procedure: 1. Clean the metal surface of thecarburetor air intake flexible duct assembly and the new P/N 0752016-10 and -11 seals with methyl ethyl keytone, methyl iso butyl keytone, acetone or lacquer thinner solvent. 2. Apply EC 847 adhesive to the carburetor air intake flexible duct assembly adapter and P/N 0752016-10 and -11 seals, allow to dry until tacky (usually between 5 and 30 minutes), and then press firmly together to insure contact (complete cure will occur within 24 hours). 3. Install the lower P/N 0752016-10 seal on the horizontal flat surface on the bottom side of the adapter flange, locating the forward edge of the seal so that it does not extend into the bend radius on the front side of the horizontal surface. 4. Pierce a hole in each P/N 0752016-11 seal to match the holes for the fasteners in the carburetor air intake flexible duct assembly adapter flange. 5. Reinstall the carburetor air intake flexible duct assembly flange to the carburetor induction air box. C. Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. Cessna Service Letter SE76-18, dated October 11, 1976, or later approved revisions, covers the subject matter of this AD. This amendment becomes effective March 1, 1977.
98-26-09: This amendment adopts a new airworthiness directive (AD), applicable to certain McDonnell Douglas Model DC 9-10, -20, -30, -40, and -50 series airplanes, and C-9 (military) airplanes, that requires a one-time visual inspection to determine if the doorstops and corners of the doorjamb of the forward passenger door have been modified, various follow-on repetitive inspections, and modification, if necessary. This amendment is prompted by reports of fatigue cracks found in the fuselage skin and doubler at the corners and doorstops of the doorjamb of the forward passenger door. The actions specified by this AD are intended to detect and correct such fatigue cracking, which could result in rapid decompression of the fuselage and consequent reduced structural integrity of the airplane.
98-26-05: This amendment adopts a new airworthiness directive (AD) that applies to all British Aerospace (Operations) Limited (British Aerospace) Model B.121 Series 1, 2, and 3 airplanes. This AD requires repetitively inspecting (using visual methods) the internal and external surfaces of the brake torque tube assemblies in the cockpit area for cracks. This AD also requires obtaining and incorporating repair procedures for any brake torque tube assembly found cracked. This AD is the result of mandatory continuing airworthiness information (MCAI) issued by the airworthiness authority for the United Kingdom. The actions specified by this AD are intended to detect and correct cracks in the brake torque tube assemblies, which could result in reduced brake efficiency with possible reduced and/or loss of airplane control.
2013-14-01: We are adopting a new airworthiness directive (AD) for certain Pilatus Aircraft Ltd. Model PC-6/B2-H4 airplanes. This AD results from mandatory continuing airworthiness information (MCAI) issued by the aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as faulty rivets installed in the airframes during production could reduce the structural integrity of the airplane. We are issuing this AD to require actions to address the unsafe condition on these products.
81-21-03: 81-21-03 BRITISH AEROSPACE CORPORATION: Amendment 39-4229. Applies to all operators of DH-114 Heron Series 1B and all series 2 airplanes which have been modified in accordance with Modification 154 or 469. To ensure wing leading edge access doors remain properly secured, accomplish the following unless already accomplished: 1. Within 350 hours time in service after the effective date of this AD, modify the wing leading edge access doors in accordance with British Aerospace Corp. Manufacturer's Technical News Sheet W.1 1, Issue 2, and Modification No. Heron 1171 (Amendment No. 1 incorporated) dated October 17, 1961, or a later FAA approved revision. 2. Alternate means of compliance may be used which provide an equivalent level of safety when approved by the Chief, Seattle Area Aircraft Certification Office, FAA, Northwest Region. 3. Airplanes may be flown to a maintenance base for repairs or replacement in accordance with FAR 21.197. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the addresses listed above. These documents may also be examined at FAA, Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective October 18, 1981.
51-04-02: 51-04-02 WARNER: Applies to All Engines Except Those With the Letter "A" Stamped After the Serial Number on the Engine Nameplate. The Letter "A" Indicates That the Modifications Hereinafter Discussed Have Already Been Accomplished. Compliance required as indicated. As a precautionary measure to preclude serious flight hazards resulting from failure of the 5/16-inch cylinder holddown studs P/N S506, the following inspections or replacements should be accomplished. A. Compliance required as soon as possible but not later than April 1, 1951 and at each 100 hours of operation thereafter. 1. Check the cylinder holddown nuts to determine that they are tightened to a torque of 180 inch- pounds desired to a 200 inch-pounds maximum. (a) If one or more of the studs is found to be broken, replace all of the studs for the particular cylinder. (b) Loose cylinder holddown nuts are a good indication that the stud may have stretched to the extent that fatigue failure may soon result. If one or more of the nuts is found loose, it is recommended that all of the studs in the cylinder be replaced. (c) When installing a cylinder with oil on the stud threads, tighten nuts to 200 inch-pounds desired to 225 inch-pounds maximum. (Warner Service Letter No. A-15 discusses procedures for inspecting and replacing 5/16-inch studs and nuts.) B. Compliance not required, but will eliminate foregoing mandatory inspections. 1. Replace cylinder stud P/N S506 and nut N-511 with 3/8-inch stud S-911 and nut N-910. 2. Replace gasket between cylinder barrel and crankcase with an "O" type ring. 3. Stamp letter "A" after serial number on engine nameplate when this modification has been completed. (Warner Service Letter No. A-17 explains the details connected with the replacement of these parts.) The Warner Aircraft Co., P. O. Box 229, Niles, Mich., which purchased the assets of the Warner Division of the Clinton Machine Co., Detroit, Mich., have a large supply of replacement part stocks that will be held in their inventory for an indefinite period. Warner Service Letter No. A-17 should be consulted before ordering these replacement parts. If either Letter A-17 or Letter A-15 is not available, copies may be obtained from the Warner Aircraft Co., Box 229, Niles, Mich.
2010-17-11R1: We are revising an existing airworthiness directive (AD) that applies to all Dowty Propellers R408/6-123-F/17 model propellers. That AD currently requires initial applications of sealant between the bus bar assembly and the backplate assembly of certain line-replaceable units, and repetitive applications of sealant on all R408/6-123-F/17 model propellers. This new AD requires the same actions and allows the use of an equivalent sealant as prescribed in revised service information. This AD was prompted by the need to add an optional terminating action to the applications of sealant. We are issuing this AD to prevent an in-flight double generator failure, which could result in reduced control of the airplane.
76-06-02: 76-06-02 BELL: Amendment 39-2554. Applies to Model 205A-1 helicopters, Serial Numbers 30001 through 30196, certified in all categories. Compliance required within the next 25 hours' time in service after the effective date of this AD, unless already accomplished. (a) Gain access to the engine fire extinguisher discharge cartridge, P/N 13083-45, mounted on the fire extinguisher agent container, as necessary, to determine if the legend "1272" is stamped on one of the hexagonal flats on the cartridge body. If this legend is found, proceed as follows: (1) Disconnect the circuit breaker for the fire extinguisher system circuit. (2) Disconnect aircraft wiring from the fire extinguisher discharge wiring. (3) Using a short length of wire (safety wire is suitable), connect the actuation pin to the ground stud. (4) Identify the 1 1/4" hexagonal nut into which the cartridge is threaded. Use an appropriate wrench to prevent this nut from turning and remove the cartridge from the nut. Leave the wire specified in item (3) attached on the cartridge terminals to prevent inadvertent detonation by static electricity. (5) Attach a connecting wire as described in step (3) above to a suitable replacement cartridge and install the cartridge into the 1 1/4" hexagonal nut from which the defective cartridge was removed. (6) Tighten the new cartridge to 90-100 inch pounds and lockwire to the 1 1/4" hexagonal nut. (7) Remove the pin-to-stud connecting wire and reconnect aircraft wiring. Close the circuit breaker. (b) If initial inspection of the cartridge reveals that the cartridge is not from the lot marked "1272," no further action is required to comply with this AD. Completion of Steps 1, 2, and 4 of Bell Helicopter Service Bulletin No. 205-75-7, dated November 5, 1975, or later approved revision, will satisfy the requirements of this AD. Equivalent methods of compliance with this airworthiness directive must be approved by the Chief, Engineering and Manufacturing Branch, Flight Standards Division, Southwest Region, Federal Aviation Administration, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment becomes effective April 26, 1976.
98-25-13: This amendment adopts a new airworthiness directive (AD) that is applicable to McCauley Propeller Systems (formerly McCauley Accessory Division, The Cessna Aircraft Company) Models 2A36C23/84B-0 and 2A36C82/84B-2 propellers. This action supersedes priority letter AD 89-26-08 that currently requires penetrant inspections for cracks in the propeller blade threaded retention area, and modifying the propeller hub to a red dye filled configuration. This action adds an explanatory note to better define the AD applicability and makes minor adjustments to compliance section language to reflect current AD practice. This amendment is prompted by reports of confusion from operators as to if the AD is applicable to their particular model propeller. The actions specified by this AD are intended to prevent possible cracks in the propeller blade threaded retention area from progressing to blade separation, which can result in loss of aircraft control.
75-14-01: 75-14-01 MESSERSCHMITT BOELKOW BLOHM Gmbh.: Amendment 39-2245 as amended by Amendment 39-2277. Applies to Messerschmitt Boelkow Blohm (MBB) Model BO-105 helicopters certificated in all categories. Compliance is required as indicated. To prevent failure in service accomplish the following: (a) Before further flight, remove flexible hose assemblies having any of the following MBB or ESPA part numbers that have blue fittings marked MS24590. Replace with flexible hose assemblies of the same MBB or ESPA part number that have silver or metal colored fittings marked 12 LN 29813. MBB PART NUMBER ESPA PART NUMBER 105 - 61795 40848 - 1290 105 - 61796 40561 - 1185 105 - 61797 40561 - 820 105 - 61798 40560 - 440 105 - 62163 40558 - 415 105 - 62165 40558 - 500 105 - 61793 40561 - 355 105 - 62161 40561 - 960 105 - 62162 40561 - 1010 105 - 62169 40848 - 820 105 - 62168 40848 - 775 105 - 61792 40562 -330 -0 105 - 61791 40562 - 1125 105 - 62166 40849 - 810 105 - 62167 40849 - 865 105 - 61799 40560 - 600 105 - 61343 40848 - 250 105 - 61344 40848 - 890 105 - 90897 40556 - 350 (Replaces 105-90898) (b) Before further flight, except that the aircraft may be flown in accordance with FAR Sec.21.197 and 21.199 to a base where the work can be performed, remove flexible hose assemblies, MBB part numbers 105-61797, -61798, and -61799 that have the marking "HN-10" behind the part/drawing number or that have "ESPA HN-10" marked on the hose surface. Replace with serviceable assemblies of the same part number that do not carry the designation "HN-10" or "ESPA HN-10". (c) Before further flight, except that the aircraft may be flown in accordance with FAR Sec. 21.197 and 21.199 to a base where the work can be performed, inspect flexible hose assemblies having the following part/drawing numbers that carry the designation "HN-10" or "ESPA HN-10",for cracks in the area where the hose extrudes from the fitting. Remove and replace in accordance with the following schedule: 105 - 61795 105 - 62162 (or D133 - 1580) 105 - 61793 105 - 62169 (or D133 - 1585) 105 - 61343 (or D133 - 1577) 105 - 61344 (or D133 - 1578) (1) If cracks are found, or prior to reaching the hose assembly service life limit if no cracks are found, remove the assembly and replace with a serviceable assembly of the same part number that does not carry designation "HN-10" or "ESPA HN-10". (2) If no cracks are found, the hose assembly may be continued in service until reaching the service life limit of 100 hours total time in service provided the assembly is reinspected at intervals not to exceed 25 hours time in service or two calendar weeks, whichever occurs first. (Messerschmitt Boelkow Blohm BO-105 Alert Bulletins 10, 11, and 12 and Service Bulletin 60-14 cover this same subject.) Amendment 39-2245 was effective upon publication in the Federal Register. This Amendment is effective upon publication in the Federal Register as to all persons except those persons, to whom it was made immediately effective by the telegram dated June 25, 1975, which contained this amendment.
90-08-03: 90-08-03 FOKKER: Amendment 39-6559. Docket No. 89-NM-238-AD. Applicability: Model F-27 series airplanes, 10102 through 10684, 10686, 10687, and 10689 through 10692, certificated in any category. Compliance: Required as indicated, unless previously accomplished. To prevent reduced structural capability of the wing due to fatigue cracks, accomplish the following: A. Perform a one-time X-ray inspection of the top skin Stringers 5 and 6 at Wing Stations 11260, 11860, 12660, and 13460, in accordance with Fokker Service Bulletin F27/57-65, dated October 4, 1989; or Revision 1, dated December 1, 1989, following the schedule below: 1. For airplanes that have accumulated less than 40,000 landings as of the effective date of this AD, inspect prior to the accumulation of 30,000 landings or within 180 days after the effective date of this AD, whichever occurs later. 2. For airplanes that have accumulated 40,000 landings or more but less than 50,000 landings asof the effective date of this AD, inspect within 90 days after the effective date of this AD. 3. For airplanes that have accumulated 50,000 landings or more as of the effective date of this AD, inspect within 30 days after the effective date of this AD. B. If cracks are found, repair prior to further flight, in accordance with Part 2 of the Accomplishment Instructions of Fokker Service Bulletin F27/57-65, dated October 4, 1989; or Revision 1, dated December 1, 1989. C. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Northwest Mountain Region. NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector (PMI), who will either concur or comment and then send it to the Manager, Standardization Branch, ANM-113. D. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Fokker Aircraft USA, Inc., 1199 N. Fairfax Street, Alexandria, Virginia 22314. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or Standardization Branch, 9010 East Marginal Way South, Seattle, Washington. This amendment (39-6559, AD 90-08-03) becomes effective on May 4, 1990.
65-13-03: 65-13-03 FAIRCHILD: Amdt. 39-85 Part 39 Federal Register June 18, 1965. Applies to Model F-27 Aircraft Serial Numbers 1 through 95. Compliance required within the next 50 hours' time in service after the effective date of this AD unless already accomplished. To prevent further malfunctions of the actuator shaft to flap gear box connecting shaft universal joints resulting in an asymmetric flap condition, accomplish the following: Modify the actuator shaft to flap gear box connecting shaft universal joints in accordance with Fairchild Service Bulletin No. 27-35, dated June 15, 1962, or later FAA-approved revision, or an equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. This directive effective July 18, 1965.
91-05-03: 91-05-03 DORNIER: Amendment 39-6902. Docket No. 90-CE-49-AD. Applicability: Models Do228-100, Do228-101, Do228-200, Do228-201, Do228-202, and Do228-212 airplanes (serial numbers as indicated in the body of the AD), certificated in any category. Compliance: Required within the next 300 hours time-in-service after the effective date of this AD, unless already accomplished. To assure the electrical bonding integrity of the affected airplanes, accomplish the following: (a) For serial numbers (S/N) 7000 through 7168 and S/N 8000 through 8190 airplanes, replace the 4 mm2 cross-sectional area bonding straps between the horizontal stabilizer and the elevator with 6 mm2 cross-sectional area bonding straps in accordance with the instructions in Dornier Service Bulletin (SB) No. SB-228-106, Revision 1, dated December 11, 1989. (b) For S/N 7000 through 7168, S/N 8000 through 8175, and S/N 8177 airplanes, visually inspect the wing front spar area around electrical connectors 56VP, 57VP, 58VP and 59VP (electrical connectors 23QXa, 24QXa if option IK04 is installed) for corrosion in accordance with the instructions in Dornier SB No. SB-228-152, Revision 1, dated February 19, 1990. If corrosion is found, prior to further flight, remove the corrosion and treat the affected area in accordance with the instructions in Dornier SB No. SB-228-152, "Accomplishment Instruction" paragraph 2.2. (c) For S/N 7000 through 7168 and S/N 8000 through 8179 airplanes, install an additional grounding strap between the wing rear spar and the fuselage in accordance with the instructions in Dornier SB No. SB-228-162, dated February 19, 1990. (d) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a location where the requirements of this AD can be accomplished. (e) An alternate method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager,Brussels Aircraft Certification Staff, Europe, Africa and Middle East Office, FAA, c/o American Embassy, 1000 Brussels, Belgium. The request should be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Brussels Aircraft Certification Office. (f) All persons affected by this directive may obtain copies of the documents referred to herein upon request to Dornier Luftfahrt GmbH, Product Support, P.O. Box 3, D-8031 Wessling, Federal Republic of Germany; Telephone (498153)-300; Facsimile (498153)-30.29.85; or may examine these documents at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64106. This amendment (39-6902, AD 91-05-03) becomes effective on March 25, 1991.
80-07-03: 80-07-03 CESSNA: Amendment 39-3723. Applies to the following Model R172K airplanes certificated in all categories: (Serial Numbers R1722000 through R1723204, R1723206 through R1723211, R1723213, R1723214, R1723216 through R1723220, R1723222 through R1723226, R1723228 through R1723238, R1723240 through R1723246, R1723248 through R1723258, R1723260 through R1723272, R1723274 through R1723312, R1723314, R1723316 through R1723319, R1723321 through R1723329, R1723331 through R1723351, R1723353 through R1723362 airplanes). COMPLIANCE: Required as indicated unless already accomplished. To preclude failure of the engine oil pressure pump drive shaft and resulting oil pressure loss caused by an improperly installed tachometer shaft connector, within the next 10 hours' time- in-service after the effective date of this AD, accomplish the following in accordance with the applicable Cessna Service Manual: A) Remove Cessna P/N 0550297-1 tachometer drive adapter assembly from the engine accessory case. B) Using a five-eighths inch crow foot wrench, measure the breakaway torque of Cessna P/N 0550296-1 tachometer drive connector. NOTE: Refer to Advisory Circular AC65-9A, Airframe & Powerplant Mechanics General Handbook, or Torque Wrench Manufacturers Instructions for formula for use of torque wrench with extensions. C) If the breakaway torque is at least 200 inch-pounds and not over 350 inch- pounds, retorque the Cessna P/N 0550296-1 tachometer drive connector to 280 to 300 inch- pounds. Reinstall the tachometer drive adapter assembly and make the prescribed entry in the aircraft maintenance record indicating compliance with this AD. D) If the breakaway torque is less than 200 inch-pounds or greater than 350 inch- pounds, prior to further flight, accomplish either of the following: 1. Replace the Teledyne Continental Motors P/N 634010 Oil Pump Drive Gear and P/N MS35756-1 Woodruff Key with a new Oil Pump Drive Gear and Woodruff Key of the same part number. (Refer to Teledyne Continental Motors Overhaul Manual for I0-360 Series Aircraft Engines for procedures.) 2. Perform a magnetic particle inspection of the Teledyne Continental Motors P/N 634010 Oil Pump Drive Gear in accordance with Section VI of the Teledyne Continental Motors Overhaul Manual for I0-360 Series Aircraft Engines. Give particular attention to the Woodruff Key slot and threaded area. If the oil pump drive gear is found cracked, replace with a new part of the same part number. Whether the oil pump drive gear is cracked or not, replace the P/N MS35756-1 Woodruff Key with a new part of the same part number. (Refer to Teledyne Continental Motors Overhaul Manual for I0-360 Series Aircraft Engines for procedures.) E) Install the Cessna P/N 0550296-1 tachometer drive connector and torque to 280 to 300 inch-pounds. Reinstall the tachometer drive adapter assembly and make the prescribed entry in the aircraft maintenance record indicatingcompliance with this AD. F) Airplanes may be flown in accordance with FAR 21.197 to a location where the maintenance required by this AD may be performed. G) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing District Office, Federal Aviation Admin., Rm 238, Terminal Bldg. No. 2299, Mid-Continent Airport, Wichita, KS 67209. This amendment becomes effective on March 31, 1980, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated March 13, 1980.
2006-01-11: The FAA adopts a new airworthiness directive (AD) for all The Cessna Aircraft Company (Cessna) Models 208 and 208B airplanes. This AD requires you to install the pilot assist handle (part number (P/N) SK208-146-2) (or FAA-approved equivalent part number) and deicing boots on the cargo pod and landing gear fairings (part number (P/N) AK208-6C) (or FAA-approved equivalent part number); and make changes to the Pilot's Operating Handbook (POH) and FAA-approved Airplane Flight Manual (AFM). This AD results from reports of several accidents involving the affected airplanes during operations in flight and in ground icing conditions. We are issuing this AD to provide a safe method to detect ice, snow, frost, or slush adhering to the upper wing (a critical surface) prior to takeoff; and to reduce drag in-flight by shedding ice on the cargo pod and landing gear fairings. Ice adhering to the upper wing surface, cargo pod, or landing gear fairings could result in a reduction in airplane performance with the consequences that the airplane cannot perform a safe takeoff or climb.
85-14-05: 85-14-05 SHORT BROTHERS LTD: Amendment 39-5088. Applies to Model SD3-30 airplanes as listed in Short Brothers Service Bulletin SD3-25-37, dated June 1984, certificated in any category. Compliance is required within 90 days after the effective date of this AD, unless previously accomplished. To assure proper operation of the crew seat harness, accomplish the following: A. Install the improved crew seat harness torsion spring assembly in accordance with Short Brothers Ltd. Service Bulletin SD3-25-37, dated June 1984. B. Alternate means of compliance which provide an acceptable level of safety may be used when approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. C. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. This amendment becomes effective August 12, 1985.
90-13-02: 90-13-02 GULFSTREAM AEROSPACE CORP: Amendment 39-6660. Docket No. 90-NM-126-AD. Final copy of Priority Letter issued June 14, 1990. Applicability: Model G-1159 (G-II) series airplanes, serial numbers as follows, certificated in any category: 001 011 021 031 040 050 061 002 012 022 033 041 051 063 003 013 023 034 043 052 065 005 014 024 035 044 053 006 015 025 037 045 055 007 017 026 038 046 056 008 018 027 039 047 057 010 019 028 049 059 020 029 Compliance: Required as indicated, unless previously accomplished. To prevent reduced structural integrity of the wing, accomplish the following: A. Within the next 10 days after the effective date of this AD, perform the following inspections: 1. Perform a visual inspection to detect corrosion or cracks in the wing rear beam/upper cap angle, wing upper aft plank, and the clothespin attach fitting, part number 1159WM20016, located at fuselage station 452. 2. Inspect the lower cavity of the clothespin attach fitting, using a 4.9 mm or smaller flexible borescope through the existing drain hole. If no drain hole exists, prior to further flight, install a drain hole in accordance with Gulfstream G-II Customer Bulletin No. 42, Section V, dated December 15, 1969. 3. Inspect the upper cavity of the clothespin attach fitting, using the borescope through the gaps where the clothespin mates with the fitting. If foam filler is present and the inspection cannot be accomplished, prior to further flight, perform the inspection in accordance with a method approved by the Manager, Atlanta Aircraft Certification Office, FAA, Central Region. 4. Inspect the wing upper aft plank adjacent to the clothespin fitting for corrosion or defects using pulse echo ultrasonic equipment (CAUTION: a machined step in this area may be misinterpreted as material loss.) B. If corrosion or cracks are found, prior to further flight, replace the affected parts or repair the corroded area in a manner approved by the Manager, Atlanta Aircraft Certification Office, FAA, Central Region. C. Within 7 days after accomplishing the inspections required by paragraph A., above, submit a written report of the inspection results to the Manager, Atlanta Aircraft Certification Office, FAA, Central Region, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia 30349. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Atlanta Aircraft Certification Office, FAA, Central Region. NOTE: The request should be submitted directly to the Manager, Atlanta Aircraft Certification Office, and a copy sent to the cognizant FAA Principal Maintenance Inspector (PMI), if appropriate. The PMI will then forward comments or concurrence to the Atlanta Aircraft Certification Office. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service document (Gulfstream G-II Customer Bulletin No. 42, dated December 15, 1969) from the manufacturer may obtain copies upon request to Gulfstream Aerospace Corporation, P. O. Box 2206, M/S D-10, Savannah, Georgia 31402-9980. These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 17900 Pacific Highway South, Seattle, Washington, or at the FAA, Central Region, Atlanta Aircraft Certification Office, Suite 210C, 1669 Phoenix Parkway, Atlanta, Georgia. This amendment (39-6660, AD 90-13-02) becomes effective on July 31, 1990, as to all persons, except those persons to whom it was made immediately effective by priority letter AD 90-13-02, issued June 14, 1990, which contained this amendment.
2013-13-03: We are adopting a new airworthiness directive (AD) for certain Airbus Model A319-112, -113, and -132 airplanes; Model A320-211, -212, -214, -231, and -232 airplanes; and Model A321-111 and -131 airplanes. This AD was prompted by a report of two fatigue cracks on the left-hand and right-hand sides of the continuity fittings at the front windshield lower framing on a Model A319 series airplane. This AD requires a high frequency eddy current (HFEC) inspection for any cracking on the left- hand and right-hand sides of the windshield central lower node continuity fittings, and repair if necessary. We are issuing this AD to detect and correct cracking of the windshield central lower node continuity fittings, which could reduce the structural integrity of the airplane.